The inspection of a compressor casing is generally conducted by measuring a gap between the tip of rotor blades and the compressor case blade path.
The rotor blades and the blade path of the compressor case are an area of critical concern both in the assembly and the operation of turbine engine technology, in particular gas turbine engine technology used in the aerospace industry. Incorrect positioning of the compressor rotor within its respective compressor case will result in dubious performance parameters leading to blade clearance variations. Accordingly, blade clearance can vary over the operating points of a gas turbine engine including during ground idle, takeoff, cruise or deceleration as well as over the cycle life of the engine. These variations can be due to a number of factors including loads on both static and rotating parts and the wear of these parts. In certain situations excessive blade clearance between the rotor blades and surrounding case can result in reduced engine efficiency. Accordingly, the performance of the gas turbine engine can be adversely affected by having a non-optimal placement of the compressor rotor within its respective compressor case.
A number of systems have been developed in an attempt to optimize compressor rotor placement and reduce blade tip clearance problems found in gas turbine engines. More specifically, off-centre positioning of the compressor rotor will increase the clearance in one area around a circumference of the compressor case but at the same time inadvertently increase the risk of a tip rub situation in opposite areas of the compressor case circumference. Accurately minimizing those blade tip clearances will ensure maximum air-mass flow is directed through the working path and not lost to “blow-by” over the tips.
U.S. Pat. Nos. 6,949,939; 5,497,101; 5,166,626; U.S. Pat. No. 5,140,494; and U.S. Pat. No. 4,071,820 disclose the use of probe members particularly adapted to be inserted and fixed in one or more positions within and around the compressor casing of gas turbine engines. Systems used for determining turbine blade clearance incorporate capacitance tip clearance systems to measure the capacitance between the fixed probe and the blade tip. Some of those systems require the full working of the turbine engine to put into affect their system. In particular, methods applied to position the compressor rotor relative to the compressor case rely on shims used in each quadrant around the compressor case circumference between the blade tips and case. However, in an attempt to optimize compressor rotor performance and subsequently reduce blade tip clearance problems, the accuracy of those systems and methods appear to be compromised by the very nature of the systems and methods used resulting in the aforementioned problems including but not limited to engine lock-up.
In order to overcome the disadvantages of the prior art, the present invention provides a rotor blade system for inspecting and optimizing the performance parameters of rotors used for example, in turbine engines, in particular gas turbine engines used in the aerospace industry.